The present invention relates to the specific field of turbine engines, and it relates more particularly to the problems that arise when mounting a combustion chamber in the casings of the turbine engine.
Conventionally, in a turbojet or a turboprop, the high-pressure turbine, and in particular its inlet nozzle (HPT nozzle), the combustion chamber, and the inner and outer coverings (or casings) of the combustion chamber are all made of metal or of metal alloy. In operation, the temperature field of the walls of the chamber and the mechanical loads that apply thereto give rise to movements at the ends of those walls. Those movements are made possible by the chamber having a flexible portion in the form of hairpins that connect together the zone for engagement with the casing (the fastener flange) and the zone of the flame tube. Nevertheless, that flexible hairpin is subjected to numerous forces that fluctuate during an operating cycle because of the thermal expansion differences that exist between the chamber and the casing. This can lead to fatigue, and in the extreme to breakage phenomena that then require the entire wall of the chamber to be replaced.
Furthermore, since the types of loading are numerous and difficult to identify accurately, it is found to be complicated to define an optimized shape when designing the part, such that it is often necessary to redesign the part in order to refine its shape until an acceptable compromise is achieved between flexibility and robustness.
Thus, in patent application FR 13/58899, the present Applicant has proposed having recourse to independent metal bands mounted between the combustion chamber and the annular coverings to take the place of conventional direct connections by means of flanges. That new mounting gives a great deal of satisfaction. Nevertheless, in certain modes of operation, it has been found that dynamic excitation by sound can lead to the hairpins of the outer band breaking.